It is known that when a boundary layer is laminar, frictional drag is greatly reduced in comparison with a case where the boundary layer is turbulent. Therefore, to reduce frictional drag in a nose-shaped object of an aircraft placed in a flow, it is desirable to suppress a laminar-turbulent transition of a boundary developed formed on a surface in order to delay a transition location as far as possible.
It is also known conversely that when the boundary layer is turbulent, separation of the boundary layer is suppressed in comparison with a case where the boundary layer is laminar, leading to a large reduction in pressure drag. Therefore, the boundary layer transition from a laminar flow to a turbulent flow is also advanced in order to reduce pressure drag caused by separation of the boundary layer on a nose-shaped object of an aircraft placed in a flow.
Adjusting a surface shape of an object in order to suppress the laminar-turbulent transition of the boundary layer and thereby delay the transition location is known as natural laminarization of boundary layer. Natural laminarization of boundary layer for a supersonic aircraft has been applied to a design for a main wing of a “scaled supersonic experimental airplane” (National Experimental Supersonic Transport; NEXST-1) developed by the Japan Aerospace Exploration Agency (JAXA), and an effect thereof has been validated by a flight experiment (see Non-Patent Document 1, for example).
Incidentally, when an angle of attack, or in other words an angle formed by an airflow and an object axis, of a nose-shaped object of an aircraft placed in a supersonic flow is zero, all flows on the object surface are aligned with a generatrix, and therefore exhibit axial symmetrically.
In this case, the boundary layer flow of the object surface is uniform in a circumferential direction, and velocity variation therein appears only in a perpendicular direction to the surface and the generatrix direction. Therefore, the laminar boundary-layer flow on the object surface in this case is known as a two-dimensional boundary layer. It is known that in a two-dimensional boundary layer, the laminar-turbulent transition is governed by Tollmien-Schlichting wave type instability.
When the angle of attack is not zero, on the other hand, a differential pressure is generated between a windward ray of symmetry and a leeward ray of symmetry, and therefore the flows on the object surface are a combination of a flow in a circumferential direction (from the windward side to the leeward side) and a flow in an axial direction. In this case, the boundary layer flow also varies in the circumferential direction, excluding the windward and leeward rays of symmetry, and is therefore known as a three-dimensional boundary layer. It is known that in a three-dimensional boundary layer, the laminar-turbulent transition is governed by cross-flow instability.
Cross-flow instability has a greater destabilization effect than Tollmien-Schlichting wave type instability, and therefore, at an identical flow velocity, the transition location advances further toward a tip when the angle of attack is not zero than when the angle of attack is zero. Advancement of the transition location is greatest in a location where a circumferential angle is approximately 60 degrees from the leeward ray of symmetry (see Non-Patent Document 2, for example).
A nose of a small sized business jet aircraft developed by HONDA MOTOR CO., Ltd. is known as a natural laminar flow nose-shaped object placed in a fluid (see Patent Document 1 and Non-Patent Document 3, for example). This small sized business jet aircraft, however, is a subsonic aircraft with a maximum cruising speed of 778 km/h (216 m/s), or in other words M=0.73, at a maximum operating altitude of 12 km (−56° C.).
Peripheral flows differ greatly between cases in which the nose-shaped object is placed in a subsonic flow of M=0.73 and a supersonic flow of at least M=1.5. The reason for this is that in a supersonic flow, compressibility is strong, and in the supersonic flow, therefore, a shock wave forms on a tip of the nose-shaped object. Moreover, a pressure distribution over the object surface differs from that of the subsonic flow.
Hence, a mechanism of natural laminarization of boundary layer on a nose-shaped object differs depending on whether the object is placed in a subsonic flow or a supersonic flow. In other words, a specific natural laminar flow nose shape for suppressing the laminar-turbulent transition in a nose-shaped object of an aircraft placed in a supersonic flow by overcoming the instability described above has not yet been found.
Patent Document 1: U.S. Pat. No. 7,093,792
Non-Patent Document 1: Naoko Tokugawa, Dong-Youn Kwak, Kenji Yoshida, Yoshine Ueda: “Transition Measurement of Natural Laminar Flow Wing on Supersonic Experimental Airplane (NEXST-1)”, Journal of Aircraft, Vol. 45, No. 5, (2008), pp. 1495-1504
Non-Patent Document 2: Yoshine Ueda, Hiroaki Ishikawa, Kenji Yoshida: “Three-Dimensional Boundary Layer Transition Analysis in Supersonic Flow Using a Navier-Stokes Code”, ICAS2004-2.8.2 (2004)
Non-Patent Document 3: “Natural Laminar Flow Wing and Nose”, Honda Aircraft Company in the U.S. [online], [retrieved Jun. 1, 2010], Internet <URL:    http://hondajet.honda.com/designinnovations/naturalLaminarF low.aspx>
As described above, there are at present no precedents whatsoever natural laminarization of boundary layer on a nose of a supersonic aircraft by adjusting the shape of the nose while taking instability into consideration.
At the Japan Aerospace Exploration Agency (JAXA), for example, a silent supersonic aircraft (known as S3TD) is being designed to test an airframe design technique for reducing a sonic boom, but natural laminarization of boundary layer has not been implemented on the nose shape thereof. Hence, the present inventor investigated respective transition characteristics of the silent supersonic aircraft (known as S3TD) and four axial symmetrical shapes numerically as a preliminary to designing a natural laminar flow nose.
It was found as a result of a numerical investigation into the transition characteristic of a silent supersonic (known as 2.5 shape S3TD) that the pressure distribution (a streamline 172 in FIG. 38) in the vicinity of the windward ray of symmetry is undulating, and in accordance therewith, a cross-flow velocity direction (a streamline 172 in FIG. 39) reverses repeatedly while an amplitude of the cross-flow velocity remains low.
Results of a numerical analysis of the silent supersonic aircraft (known as S3TD) show that when the cross flow reverses repeatedly without developing in a single direction, amplification of the amplitude is suppressed. In other words, when the pressure distribution over the nose surface undulates, the cross-flow instability can be suppressed favorably such that the boundary layer transition on the nose surface can be delayed, or in other words the transition location can be delayed.
Further, it is evident from a pressure distribution shown by a streamline 116 in FIG. 38 that even through the pressure distribution undulates similarly to the streamline 172, the laminar-turbulent transition location (the boundary layer transition location) is closer to the tip than the streamline 172 (an N factor in FIG. 39).
Hence, by appropriately modifying the phase of the undulating pressure distribution, a natural laminar flow effect for reducing frictional drag by delaying the transition location and a separation suppression effect for suppressing separation by advancing the transition can be realized selectively on the nose surface.
In other words, with an undulating pressure distribution, both a natural laminar flow effect and a separation suppression effect of separation are obtained. It was therefore found that an undulating pressure distribution on the surface is extremely important for natural laminarization of boundary layer and suppressing separation on the nose.